Two stage rocket combustor

ABSTRACT

A method for operating a rocket engine by injecting fuel and oxidizer into an elongated combustion chamber in two flows, a core flow where the fuel and oxidizer are intimately mixed and immediately combusted and a peripheral curtain flow which surrounds the core flow and which is in contact with the combustion chamber wall to cool it and limit the heat transfer from the wall to the injector to prevent vapor locks in the injector. To prevent decomposed or partially combusted propellant products from chemically reacting with the chamber wall no mixing of fuel and oxidizer takes place in the curtain flow. The curtain flow is deflected radially inward into the core flow, before decomposed or partially combusted products can come into contact with the wall, into the core flow to fully combust the curtain flow out of contact with the wall. The rocket engine is defined by serially arranged first and second combustion chambers and an injector constructed to form the core and curtain flows. A ring plate projects radially inward at the downstream end of the first chamber and deflects the curtain flow into the coreflow.

This is a division of application Ser. No. 172,622, filed Mar. 24, 1988,now U.S. Pat. No. 4,882,904.

BACKGROUND OF THE INVENTION

A conventional liquid fuel rocket engine has a propellant (fuel andoxidizer) injector and a nozzle shaped combustion chamber into which thepropellant is injected and where its constituents are mixed and oxidizedor combusted.

There are a number of proven fuel and oxidizer combinations, each withspecific performance characteristics. They differ in the amount ofenergy released when they combust and in the thermodynamics of theirreactions. By the nature of their reactions all the propellantcombinations are exothermic, that is, once the fuel and oxidizer aremixed, the reaction is initiated and energy is released. Once thereaction has been initiated, the release of energy from the propellantcombination will drive the reaction.

The efficiency of the reaction, that is, the degree to which bothpropellant components are completely reacted, is largely dependent onthe thoroughness of the mixing of the two components. Incomplete mixingresults in unreacted fuel or oxidizer and a corresponding loss inefficiency.

The propellant injector of the rocket engine implements the mixing ofthe fuel and oxidizer components of the propellant. A typical injectorcan have from several to several thousand orifices through which thefuel and oxidizer are introduced into the combustion chamber. Theorifices direct the fuel and the oxidizer so that they form spray fanswhich mix and commence combustion at preselected points within thecombustion chamber, typically immediately down stream of the injector.Numerous orifice arrangements which form fuel only, oxidizer only orbipropellant fans exist.

The combustion of rocket propellant generates gas temperatures whichgenerally exceeds the melting temperature of most known materials usedin the construction of chamber walls. Without cooling the chamber wallswould deteriorate and ultimately melt.

Further, the heating of the combustion chamber walls can lead to anoverheating of the injector as a result of heat transfer between themvia their interface. This can result in a vapor lock in the injector,and engine failure.

There are two main approaches to cooling the combustion chamber walls ofa rocket engine. Regenerative cooling circulates one or both of thepropellant components through the walls of the chamber. The propellantcomponent acts as a coolant and carries away the heat which iseventually returned to the combination gases. The approach has onlylimited utility. It can not provide sufficient cooling for small enginesbecause their propellant flow is too low and this approach may not besuited for use with large, high-pressure engines for other reasons.

The other often practiced approach employs film cooling in which theorifice pattern of the propellant injector generates two propellantflows, a central core flow and a peripheral or curtain flow whichsurrounds or envelopes the core flow. In the core flow the propellant iswell mixed and combusted in the core which is some distance radiallyinward of the chamber wall. The cooler curtain is formed by unmixed andtherefore uncombusted propellant directed by injector orifices towardthe chamber walls. The unmixed propellant forms a cool gas film orcurtain over the chamber wall which separates it from the very hot coreflow. The film absorbs heat by evaporation of the small fuel and/oroxidizer droplets ejected by the injector and thus insulates thecombustion chamber wall from the heat of the core flow. This method canbe used with most rocket engines, but the film of uncombined propellantin contact with the combustion chamber wall is disadvantageous.

The propellant film evaporates and decomposes from its exposure to theheat. Decomposition products react with unmixed propellant to create avariety of aggressive chemical species which can chemically react withtypical chamber wall materials such as copper, nickel, platinum,iridium, gold, rhenium and columbium. This corrodes and deteriorates thechamber wall and can lead to its failure.

Further, the propellant film can undergo spontaneous thermaldecomposition, resulting in transient species which can be the source ofadditional chemical attack on the combustion chamber wall.

Partially mixed propellant can result in localized concentrations offuel and oxidizer existing side by side. The boundaries between theseconcentrations create an environment where a spectrum of combustionchemistry species are generated, including many nonequilibrium speciesnot normally found in other combustion devices. When these species comein contact with the combustion chamber wall, they cause a conditionknown as a streaking, blanching or scalloping of the walls. They canalso attack the injector orifices, resulting in what is commonly calledbell-mouthing.

A further disadvantage of prior art film cooling is that it decreasesthe efficiency of the rocket engine. The quantity of the propellant usedfor cooling is significant and, to a substantial extent, its use as acoolant causes it to be lost to the system as propellant. This canresult in an appreciable loss of efficiency.

SUMMARY OF THE INVENTION

The present invention provides an improved film cooling method forkeeping the combustion chamber wall of a rocket engine cooled. Briefly,the chamber is divided into upstream and downstream sections. Theinjector and the chamber are constructed so that only unmixed andnondecomposed propellant comes in contact with the upstream section ofthe chamber and only fully combusted propellant contacts the downstreamsection. The chamber is further constructed so that all propellant iscombusted in the core of the chamber and substantially no propellant iscombusted along the chamber walls.

A principal advantage of the invention is that it inhibits chemicalreactions between various gas species and the combustion chamber wall.The integrity of the combustion chamber wall is thereby preserved.

Another advantage of the invention is that it enhances the efficiency ofthe propellant combustion. This is accomplished by deflecting thecurtain flow from the chamber wall radially inwardly into the core flowso that the fuel and/or oxidizer which makes up the curtain flow isintimately mixed and fully combusted. The heretofore encountered loss offuel and/or oxidizer in the curtain flow due to non or incompletecombustion is thereby eliminated. Further, the decomposition andcombustion of the curtain flow, which creates the above-discussedchamber wall corroding chemical species, occurs in the core flow, spacedaway from the chamber wall so that it can not deteriorate the wall.

A radially inwardly oriented, annular baffle ring protrudes from theinside of the chamber wall and deflects the curtain flow into the coreflow. As the curtain flow propagates from the injector to baffle ring,heat from the core flow and the mixture of the core flow with thecurtain flow at their boundaries causes the gradual decomposition andpartial combustion of the propellant constituent which makes up thecurtain flow. This increases from the boundary of the two flows towardthe chamber wall as the distance from the injector increases. The baffleis positioned so that the curtain flow is deflected inwardly beforedecomposition and/or partially combusted products in the curtain flow(hereafter collectively referred to as "decomposition products") reachthe periphery thereof, i.e., before such products come into contact withthe chamber wall.

Thus, the curtain flow insulates the chamber wall from the injector tothe baffle ring from the hot core combustion decomposition products.Chamber wall corrosion and deterioration is thereby minimized orprevented. The cooler curtain flow also permits a greater choice in theselection of materials in the injector to baffle region. It is oftendesirable to employ materials having low thermal conductivity to furtherinhibit heat flow from the hot region down stream of the baffle to theinjector.

The portion of the chamber downstream of the baffle ring comes intocontact with substantially only fully combusted products. i.e.,propellant which has already been combusted in the core flow. Thedecomposition products of the curtain flow are intimately intermixed andcombusted within the core flow. Since only combusted propellant comesinto contact with the downstream chamber wall the wall can beconstructed of high temperature, thermally highly conductive materialswhich may be subject to chemical attack by decomposition products butnot by fully combusted propellant. This greatly facilitates theconstruction of the high temperature downstream chamber section forcooling by radiation which requires both high thermal conductivity and athin wall section for a high rate of heat transfer.

As the foregoing has demonstrated, the present invention effects anefficient cooling of the upstream chamber wall section immediatelydownstream of the injector and further inhibits heat flow through theselection of low thermal conductivity materials. The transfer of heatthrough the interface between the chamber wall and the injector canthereby be limited and controlled to prevent vapor locks. The presentinvention further isolates the upstream chamber wall from thedecomposition products, thereby preventing them from chemicallyattacking and deteriorating or corroding the wall. In the preferredembodiment of the invention, this is accomplished by construction theinjector so that only one of the propellant constituents, e.g., eitherfuel or oxidizer, forms the curtain flow. It is possible, however, topractice the present invention by directing one or more alternatingzones of the propellant constituents into the curtain flow. At the axialinterfaces between the zones, some fuel and oxidizer can mix and somedecomposition product may form and chemically react with the chamberwall. Nevertheless, overall chamber wall deterioration is still greatlyreduced as compared to prior art curtain flow systems because thecurtain flow is deflected into the core stream before the decompositionproducts within the zones can reach the chamber wall. In addition,combustion efficiency is greatly improved for the reasons discussedearlier.

BRIEF DESCRIPTION OF THE DRAWING

The drawing is a side elevational view, in cross section, through arocket engine having an injector, a combustion chamber and a nozzle andit schematically illustrates the present invention.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring to the drawing, a rocket combustor chamber 2, constructed inaccordance with the present invention generally comprises an injector 4,a combustion chamber 6 for propellant injected by the injector, and anozzle 8 through which combusted gasses escape from the chamber, therebygenerating the desired rocket thrust.

The injector includes at least several and typically a multiplicity oforifices 10 through which the fuel and oxidizer constituents of thepropellant are ejected into the combustion chamber. Fuel and oxidizerare conventionally supplied to the upstream end of the orifices.

The orifices are arranged so the fuel and oxidizer ejected from theirrespective downstream ends form predefined flow patterns which generatea propellant core flow 12 that propagates in a downstream directionthrough the combustion chamber 6 and nozzle 8. The core flow supplyingorifices are oriented so that fuel and oxidizer impinge and intimatelymix a short distance downstream of the injector to initiate and completecombustion in the core stream as soon as possible.

Orifices 10a are arranged along the radially outer portion of theinjector. They are oriented to form a curtain flow 16 of preferablyeither the fuel or the oxidizer constituent of the propellant. Thecurtain flow is annular in cross-section and envelopes the core flowover its entire periphery.

The combustion chamber 6 is constructed of first and second, upstreamand downstream cylindrical combustion chamber walls 18, 20. The wallsare joined, for example with a flange 22 suitably secured, e.g., bolted,welded or the like to the chamber walls, and they define a primarycombustion chamber 24 immediately downstream of injector 4 and asecondary combustion chamber 26 located generally between the primarychamber and nozzle 8. The upstream end of the first cylindrical wall mayinclude a connector flange 28 for securing it to the injector, forming aheat transferring interface 30 between them.

In the preferred embodiment, the downstream end 32 of the firstcylindrical chamber wall 18 is defined by an annular baffle ring 32which projects radially inwardly from the first chamber wall. In theillustrated embodiment, the baffle ring is integrally constructed withthe first chamber wall. Alternatively, it may be secured thereto, as bywelding, for example. Moreover, the baffle ring may be positionedslightly upstream of the downstream end of the first chamber wall (notillustrated). In either event, the distance between the injector and anupstream facing deflection surface 36 of baffle ring 34 is dimentionedas is further described below.

In use, propellant flows through the injector 4 into combustion chamber6 as described above to form a high-temperature core flow in which thepropellant is combusted from a point just downstream of the injector.The curtain flow 16 surrounds the core flow and limits the amount ofheat that is transferred from the core flow to the first chamber wall18. For a given design, the radial thickness of the curtain flow isselected to limit the heat transfer from the first chamber wall viainterface 30 to the injector and prevent the formation of vapor locks inthe injector, or in the supply conduits (not shown) upstream thereof.

The curtain flow is preferably a homogeneous flow of one of thepropellant constituents. i.e., the fuel or the oxidizer. It alsoisolates the first chamber wall from coming into contact, and therebybeing attacked by chemically reactive species which initially form at aboundary 38 between the core and curtain flows as the fuel or oxidizerin the curtain flow thermally decompose and become partially combustedunder the heat of the core flow. This thermal decomposition and partialcombustion of the curtain flow propagates radially outward from theboundary 38 as the curtain flow travels in a downstream direction. Thedistance between the injector 4 and deflection surface 36 of baffle ring34 is selected so that the curtain flow is deflected radially inwardlyinto the core zone before the heat from the core flow, and intermixingbetween the core flow and the curtain flow at their boundaries, causesthe occurrence of chemically reactive species in the peripheral layer ofthe curtain flow, i.e., before such species can come into contact withthe first chamber wall.

The inward deflection of the curtain flow turbulently mixes it with thecore flow so as to bring substantially all constituents of the curtainflow into contact with uncombusted or partially propellant combustedconstituents in the core flow. This effects the substantially completecombustion of all propellant and maximizes the efficiency of thecombustion process by avoiding the discharge of incompletely combustedpropellant from the rocket engine as was the case with prior art rocketengines employing film cooling.

As a result of the inward deflection of the curtain flow, the secondchamber wall comes in contact with primarily only fully combustedpropellant which does not chemically react with many materials, such asmolybdenum copper, nickel, gold, platinum, iridium, rhenium etc, wellsuited for the cooling of the wall. Thus, the material of the secondchamber wall can be chosen so that it provides the desired high rate ofheat transmission through the wall without having to compromise thisobjective with an increase of the chamber wall thickness, to compensatefor chamber wall corrosion and deterioration.

As the foregoing demonstrates, the present invention conceptuallydivides the combustion chamber of a rocket engine into coaxial primaryand secondary chambers. The first chamber wall forming the primarycombustion chamber comes into contact with only the curtain flow andtherefore, with only uncombusted propellant. Conversely, the secondchamber wall defining the secondary combustion chamber comes intocontact with only substantially fully combusted propellant. The rocketengine is thereby effectively cooled, the combustion chamber isprotected against chemical attack from the decomposing and combustingpropellant, it has a significantly improved combustion efficiency ascompared to similarly cooled, though not chemically protected prior artrocket engines, and the respective chamber walls can be constructed ofdifferent materials, each best suited to for its particular function.

I claim:
 1. A method of operating a rocket engine by combusting a fuelwith an oxidizer, the method comprising the steps of:a. forming firstand second, axially joined peripheral walls defining contiguous firstand second combustion chambers; b. injecting proportionate amounts offuel and oxidizer into the first chamber from one end thereof oppositethe second chamber in a pattern to form a core flow which is spacedradially inward of the first and second walls and a curtain flowenveloping the core flow and in contact with the first wall from aboutthe one end to about the other end of the first wall; c. proportioningthe injected fuel and oxidizer so that the core flow forms a hightemperature flow in which fuel and oxidizer are intimately mixed andundergo combustion and so that the curtain flow forms a relatively lowtemperature flow in which substantially no fuel and oxidizer are mixed;d. directing the curtain flow in a generally radially inward directionaway from the first wall at a distance from the one end of the firstchamber where, in cross-section, a radially outermost portion of thecore flow in contact with the first wall comprises substantially nodecomposed or partially combusted fuel and oxidizer; e. substantiallyfully combusting the curtain flow radially inward of the second wall inthe core flow as the core flow propagates through the second combustionchamber.
 2. A method according to claim 1 wherein the step ofproportioning includes the step of limiting the curtain flow tosubstantially only one of the fuel and the oxidizer.